The statements in this section merely provide background information related to the present disclosure and may not constitute prior art.
Spacecraft require safe, lightweight, affordable heat shields for protecting the spacecraft and its occupants during re-entry of the spacecraft into the Earth's atmosphere. Traditionally, the weight of the heat shield has been an important factor. The higher the weight of the heat shield panels used the lower the payload that the spacecraft will be able to carry. Previously manufactured heat shields have typically been made from mixtures of epoxy-phenolic resins with fillers, from phenolic resins with fillers, from carbon-carbon composites with backside insulation, from quartz-phenolic composites with backside insulation, or from Phenolic Infiltrated Carbon Ablator (PICA) material. However, existing solutions can often add significant weight to a spacecraft.
The cost of manufacturing previously developed heat shields for a spacecraft has also been an important concern for designers. Traditionally, the high cost of manufacturing heat shields for spacecraft has contributed significantly to the overall cost of manufacture for a spacecraft. The manufacturing of previously developed heat shields has also often required complex manufacturing processes.
When manufacturing heat shields using a curable ablative material, in some limited instances there may be a tendency for the ablative material to crack during the cure process. This can be due to shrinkage of the ablative material during a curing process. It may also be due to coefficient of thermal expansion differences between the ablative material and other substrate materials that help to form the heat shield. In the event any small or large cracks develop, then such cracks would need to be repaired subsequent to completion of the curing process and any following cool-down period.
Options for attempting to prevent cracks from developing could involve making relief cuts in one of more layers of material of the heat shield. This may reduce or eliminate the possibility of one or more large cracks developing in the ablative material during the curing process. However, such an approach would still require post-processing work to fill and repair the intentionally made relief cuts.
Another option for attempting to prevent cracks in the ablative material during the cure process would be to add chemicals to the ablative material that would attempt to limit shrinkage of the ablative material during the cure process. Obviously, this requires modification of the chemical structure of the ablative material, and such may not be desirable because of the risk that the additional chemicals might interfere with the overall effectiveness of the ablative material in acting as a heat shield.
Still another option might be to cure the ablative material of the heat shield to a less mature state. However, this option might not enable the ablative material to provide the same level of ablative performance. This option may also introduce potential problems relating to “heat soak” which may occur in situ. By “heat soak” it is meant the thermal exposure that the material would encounter during flight or mission conditions. During a normal mission of a spacecraft, each component of the spacecraft requiring a heat shield experiences different thermal environments. With the less mature cure option, if the thermal conditioning experienced by the heat shield during a mission prior to the point at which the ablative heat shield needs to function is sufficient to advance the cure of the ablative material, then it may potentially cause the same cracking that the less mature state was used to avoid. As such, this option makes the expected thermal environment that the components are expected to experience prior to the point at which ablation is required a more important consideration.
Still another option might be to cure the ablative material separately from the other components of the heat shield to which the ablative material will ultimately be secured, and then perform a secondary bonding operation to permanently attach the ablative material layer to one of the other material layers of the heat shield. However, secondary bonding would significantly increase the overall complexity of the heat shield by requiring gaps, joints and surface machining of the bond surface on the ablative material.